The invention concerns a sheet entity, in particular for use as a skin plate for an aircraft fuselage and also an aircraft fuselage with at least one sheet entity of this type.
Civil aircraft are conventionally manufactured with a fuselage made from metal skin panels. In recent times, however, composite materials such as aluminum laminates with a multiplicity of sandwich-type bonded aluminum layers and glass fiber layers (GLARE®) have been deployed for purposes of optimizing the fuselage. A current example is the Airbus A380 wide-body aircraft, whose outer fuselage shell comprises an aluminum laminate of this type. Likewise carbon fiber-reinforced plastic laminates (CFRPs or CFCs) are used for purposes of optimizing the fuselage. Both the plastic laminate and also the aluminum laminate are distinguished by a lower weight compared with a conventional metal skin panel, wherein the aluminum laminate in particular has a higher fatigue strength and a lower crack propagation rate than a conventional metal skin panel. Furthermore the aluminum laminate is comparatively robust with regard to mechanical damage and has high acoustic insulation and thermal insulation properties. It is therefore an endeavor to manufacture the fuselage from a multiplicity of skin panels with different types of laminate. However, the connection of the plastic laminates with the aluminum laminates is problematical, inasmuch as in the event of contact of the aluminum layers with the carbon fiber layers, the aluminum layers tend to corrode, such that any direct contact between these two materials must be avoided at all costs. For the solution of this problem it is proposed by the applicant in the German patent application DE 10 2007 046 478 A1 to insert a solid titanium profile into the respective longitudinal seam region of the carbon laminates and to insert a solid aluminum profile into the respective longitudinal seam region of the aluminum laminates; these profiles are then connected with one another by means of friction welding. As a result of the solid metallic longitudinal seam regions, however, the loads and stresses in the laminates are not optimally distributed over the fuselage, so that stress peaks can occur in the region of the inserts. Furthermore the mechanical properties of the laminates alter in the edge region. In addition an insert of this type is relatively cost-intensive to manufacture.